The invention relates to the field of composites and to their application in particular in aeronautical turbomachines.
In the aeronautical field in particular, a constant objective is to produce parts of great mechanical strength in an unfavorable environment for a minimal weight and size. Thus, certain parts are produced from a ceramic matrix composite (CMC), such as for example an SiC/SiC composite produced for applications of long duration at high temperatures. Such a composite consists of a multidirectional reinforcement made of silicon carbide and a matrix, also made of silicon carbide. The latter gives the composite exceptional thermomechanical fatigue resistance, in an oxidizing environment, at temperatures possibly up to 1400° C. These composites are currently used for the manufacture of aircraft engine parts, namely nozzle flaps, combustion chamber and reheat system. It is the fibers that take up the loads, the matrix providing a function of binder for the rest of the part and of protecting and isolating the fibers, which must not come into contact with the oxidizing atmosphere.
The difficulties of integrating a composite, in particular a CMC, in any environment are connected with its manufacturing tolerances: the geometrical envelope of this composite may vary up to ±0.55 mm in thickness.
When such a part is to be assembled with another part made of a CMC or a metal, the practice hitherto has been to use the technique of matching. However, with this technique it is not possible to dimension the parts separately —it is the assembly that is dimensioned. Used in development, this method cannot be economically transposed to mass production. In particular, matching eliminates interchangeability of the parts, since no unique definition exists. In addition, matching is an expensive method both in terms of production and after-sales service for spares.